Field of the Invention
The present invention relates to turbine wheel blades in a turbine engine such as an airplane turboprop or turbojet.
Description of the Related Art
Conventionally, a turbine in a turbine engine comprises alternating annular rows of stationary vanes and of rotor wheels, each wheel having a plurality of radial blades mounted at the periphery of a rotor disk. Each blade has a root connected via a tang to a platform from which there extends an airfoil. The blade roots are engaged with little clearance in substantially axial slots in the periphery of the disk, which slots are regularly distributed around the axis of the disk and define splines or teeth between them. The blades are held radially in the slots by co-operating shapes, the blade roots having a cross-section of dovetail or analogous shape, for example. In the assembled position, the platforms of the blades are arranged circumferentially side by side and they surround the teeth of the disk.
It is known for the spaces that are situated radially inside the platforms and defined circumferentially between two consecutive tangs to be partitioned in the axial direction in order to limit the flow of hot air coming from the main flow passage and going towards the disk.
For this purpose, upstream and/or downstream radial walls may connect the upstream and/or downstream edges to the blade root. It is also possible to combine the above-mentioned radial walls with the use of annular plates mounted on the upstream and/or downstream faces of the disk carrying the rotor blades. With radial walls, the circumferential edges of the blade platforms and the circumferential edges of the radial walls of two adjacent blades are spaced apart circumferentially by the small amounts of clearance that are needed for assembling the blades on the disk. When a plate is mounted on an upstream or downstream face of the disk, it may for example be engaged at its radially inner periphery in an annular groove of the disk, and at its radially outer periphery in an annular groove formed radially inside the platforms that are arranged end to end. Axial clearance may also exist between the plates and the platforms of the blades.
In operation, it is found that the hot gas from the combustion chamber can flow into the cavities situated under the platforms through the clearance between the facing circumferential edges of the platforms, thereby leading to the disk being heated and possibly being damaged.
In order to reduce the heating of the teeth of the disk, it is known (from FR 2 972 759) to mount a sealing member in each inter-blade cavity, which sealing members are urged radially outwards in operation by centrifugal force and come to press radially against the inside faces of circumferentially facing platforms in order to limit leakage of hot gas therebetween. Such sealing members may also serve to damp the vibration to which the blades are subjected in operation.
Nevertheless, in operation, it is found that hot gas from the main passage can reach the insides of the cavities by passing around the zones where the above-mentioned sealing members press against the inside faces of the platforms. Furthermore, that type of part is subjected to a large amount of wear, thereby correspondingly reducing its effectiveness and its lifetime.
In order to reduce the temperature in the inter-blade cavities, cooling air is taken upstream in the turbine engine from a low or high pressure compressor and is conveyed via ducts from a zone where the disk is attached to another disk into the inter-blade cavities so as to limit the increase of temperature therein and limit the heating of the periphery of the disk.
When the rotor wheel constitutes the first wheel that the upstream end or the last wheel at the downstream end of the turbine rotor, an annular gap is formed between the upstream first wheel and a stator element upstream from the first wheel or between the downstream last wheel and a stator element downstream from the last wheel. This annular gap thus provides direct communication between the main passage for combustion gas flow and internal elements of the turbine engine.
In order to avoid hot air being introduced to the inside of the turbine engine, it is then necessary to prevent the hot gas from the main passage flowing through such an annular gap. For this purpose, a fraction of the cooling air taken from the compressor is directed radially outwards through the annular gap, thereby compensating the pressure of the hot air stream in the main passage and keeping the hot air in said passage. Nevertheless, that requires an additional amount of cooling air to be taken off in order to obtain a flow rate of the cooling air that is great enough to prevent hot air from flowing towards the inside of the turbine.
In general, the cooling air taken from the compressor reduces the efficiency of the compressor and consequently reduces the performance of the turbine engine.